Cast gas turbine blade through which coolant flows, together with appliance and method for manufacturing a distribution space of the gas turbine blade

ABSTRACT

A cast gas turbine blade through which coolant flows, includes a blade root inserted into a disk of the gas turbine; a plurality of supply ducts; and a distribution space. Coolant is fed to the supply ducts through. a feed duct of the disk, the feed duct communicating with the supply duct via the distribution space. Flow and manufacture are optimized by a cast distribution space being present which has rounded or beveled inlet openings for the supply ducts and which is manufactured by means of a one-piece casting core.

This application is the national phase under 35 U.S.C. §371 of PCTInternational Application No. PCT/EP00/02606 which has an Internationalfiling date of Mar. 23, 2000, which designated the United States ofAmerica, the entire contents of which are hereby incorporated byreference.

FIELD OF THE INVENTION

The invention relates to a cast gas turbine blade/vane through whichcoolant flows, in particular a gas turbine rotor blade with a bladeroot, which is inserted into a rotatable disk of the gas turbine andwhich has a plurality of supply ducts for an internal cooling system,and a distribution space. Preferably, it is possible to feed coolant tothe supply ducts by way of a feed duct of the disk, wherein the feedduct communicates with the supply ducts via the distribution space. Theinvention also preferably relates to an appliance, for casting a gasturbine blade/vane, having a casting core which has core ribs formingthe supply ducts. Finally, the invention also preferably relates to amethod of manufacturing a cast gas turbine blade/vane.

BACKGROUND OF THE INVENTION

From U.S. Pat. No. 4,344,738, a gas turbine blade/vane is known which isinserted by means of a blade root into a transverse groove of arotatable disk of the gas turbine, the disk having a feed duct forsupplying the gas turbine with coolant. Below the blade/vane root, thefeed duct opens into the disk transverse groove intended foraccommodating the blade/vane root. Supply ducts, through which thecoolant is fed into the internal cooling system, emerge from the bladeroot. The supply ducts have, in the main, inlet openings with edges.

U.S. Pat. No. 4,992,026 reveals a gas turbine blade/vane through which acoolant flows and which has an internal cooling system, the coolantbeing introduced by feed ducts into the blade root and fed throughsupply ducts into the internal cooling system. At their transitions fromthe blade root, the supply ducts have right-angle edges.

The object of the internal cooling of the gas turbine blade/vane is toprevent severe heating of the blade/vane material, which occurs due tohigh operating temperatures and can lead to serious damage. For thispurpose, it is necessary that the cooling medium should, withoutdifficulty, reach the gas turbine blade/vane parts which, in particular,are remote from the inlet flow region and are exposed to the most severeeffects. In the vicinity of the supply duct inlets which, however, havepractically no cooling requirement, dead zones occur. Further, the flowsdepart greatly from the ideal laminar process in the case of inlets,which are greatly constructed with edges, to the supply ducts leading tothe internal cooling system, such as is published in U.S. Pat. No.4,344,738 or U.S. Pat. No. 4,992,026, for example. This, for example,involves an increased danger of the formation of deposits and, inparticular, a high flow resistance. It is only possible to force thecooling medium through the supply passages by means of an increasedpressure and this is frequently impossible to a sufficient extent.

A further possibility for forming the supply duct region in the lowerblade root consists in providing a so-called distribution space fromwhich the supply ducts for the internal cooling system emerge and whichis supplied with coolant by the feed duct in the disk. Essentially, thedistribution space should serve to provide a reliable and uniformdistribution of the coolant to the supply ducts and it is onlypermissible for small coolant losses to occur. In the usual castingprocess, this distribution space generally has a rectangularconfiguration and has, in particular, right-angle transitions betweenthe supply ducts and distribution space. Due to the construction withedges of the inlets to the supply ducts, strong flow eddies occur which,in principle, ensure good cooling of the regions over which flow occurs.Because, however, the distribution space is in the blade/vane root, itis not subject to severe heating effects and has, therefore, only asmall cooling requirement.

This state of affairs can be improved by mechanical reworking, after thecasting process, of the supply duct inlets from the distribution space.Because of the geometry of the blade root and the properties of theblade/vane material, however, this must generally take place manuallyand is, in consequence, very labor-intensive. Furthermore, thisprocedure does not ensure that all the supply ducts of a gas turbineblade/vane have the desired shape and that all the gas turbineblades/vanes of a type have the same flow resistance. This, however,would be necessary for a calculation in advance of the flow propertieswhich would satisfy the high quality requirements and would be necessaryfor optimum utilization of the cooling medium.

SUMMARY OF THE INVENTION

An object of the invention is therefore to provide a cast gas turbineblade/vane, in particular a gas turbine rotor blade, through whichcoolant flows. It further preferably has transitions from thedistribution space to the supply ducts which are optimized in terms offlow, i.e. which has low flow resistance at the outlet openings from thedistribution space. An object is to be able to manufacture thedistribution space and the internal cooling system in a singlemanufacturing process, the casting process. A further object of theinvention includes providing an appliance and a method for manufacturingsuch a cast gas turbine blade/vane, through which coolant flows, with acorresponding distribution space.

An object directed toward a cast gas turbine blade/vane through whichcoolant flows is achieved by a cast distribution space being presentwhich has rounded or beveled inlet openings to the supply ducts, forexample.

The rounded or beveled inlet openings to the supply ducts, which areadjacent to the distribution space, ensure that the flow resistance tothe cooling medium is minimized, particularly in the transition regionbetween the distribution space and the supply ducts. The cooling mediumflow remains substantially laminar. The coolant can therefore—given anappropriate edge-free transition solution from the feed duct to thedistribution space—flow almost unhindered into the distribution spaceand out of it through the supply ducts. In this way, it reaches theinternal cooling system rapidly and with low losses, which leads to agreatly increased service life, particularly in the case of the hot andcoolant-intensive regions of the gas turbine blade/vane, for example theleading edge region. The coolant supplied is thus utilized better.

It is no longer necessary for the medium supplied through the feed ductof the disk to be guided round two 90° angles into the internal coolingsystem. On the contrary, it is fed directly to the internal coolingsystem in a smooth, continuous flow motion. No cavitation, in which thecoolant is at rest as in dead zones, occurs while the cooling mediumflows round. Because of the rounding or beveling of the inlet openings,the cooling medium supplied is only eddied to a very slight extent.

The inlet openings to the supply ducts abut directly onto thedistribution space and are generated with it during one manufacturingprocess. The rounding or beveling is shaped by the casting process in areproducible manner. In this way, a series of gas turbine blades/vanescan have the same, predetermined sizes and dimensions for the inletopenings and the distribution space. This provides the basis for areliable determination in advance of the coolant requirement and/or thecoolant function. This is particularly important for ensuring that evenremote parts of the gas turbine rotor blades are reliably cooled andthat, therefore, the wear due to overheating is minimized.

Due to the present invention, the coolant has already been introducedthrough the distribution space into the supply ducts at a low pressurebecause of the low flow resistance and it therefore escapes to only asmall extent through the intermediate space between the blade root androtating gas turbine disk. By this means, the coolant losses areminimized and the coolant is utilized in an optimum manner.

Because the distribution space has a configuration rounded in the mannerof an ellipsoid, the cooling air can be fed particularly advantageouslyto the supply ducts. In this arrangement, the distribution space ispreferably configured in the form of a semi-ellipsoid. Its base areaalso corresponds to the maximum cross section of the ellipsoid and, inthe case of a gas turbine blade/vane inserted in a disk groove, isbounded by the disk. The side surfaces of the semi-ellipsoid, and alsothe transitions between the side surfaces, have a rounded configuration.This simple geometry can be easily manufactured and reliably preventsthe formation of dead zones in which the coolant introduced comes torest. Due to the absence of edges, only slight eddying, which leads tonegligible flow losses, occurs on the walls of the distribution space.The ellipsoid-type shape makes it possible to direct the coolant supplyin a specific manner to the supply ducts adjacent to various regions ofthe ellipsoid.

A further optimization of the coolant flow is achieved by the rounded orbeveled inlet openings meeting one another or being adjacent to oneanother in a manner which optimizes the flow. Optimized in terms of flowmeans that the flow deflections necessary due to the position of twoinlet openings relative to one another or due to the position of thedistribution space and an inlet opening relative to one another takeplace with the smallest possible amount of flow eddying. In particular,this takes place because the edges, which occur due to the meeting ofthe respective curvatures of the inlet openings, are in turn roundedoff. The optimization of the shape preventing flow eddying is producedin the casting process by the employment of the rounded, one-piece castcore in a manner which can be individually matched without reworking tothe requirements set for a certain type of gas turbine.

A predetermined supply of coolant can be easily adjusted by the crosssection of the feed duct and the local changes to the cross sections ofthe distribution space being matched to the cross sections of the outletopenings located downstream. The cross-sectional changes to the heightand width of the distribution space correspond, for example, to theshape of a semi-ellipsoid. The transitions between the inlet openings oraround the inlet openings are designated as transition cross sections.The rounding or beveling of the inlet openings produces a larger inletopening cross section directly at the distribution space and this crosssection is then reduced again on transition to the supply duct. The feedduct has an essentially constant cross section but it can also berounded or beveled in order to improve the flow properties, thusincreasing the cross section in the direction toward the distributionspace. The cross sections described are matched to one another, i.e.predetermined cross-sectional relationships are taken into account inthe matching of the coolant supply. This is necessary if, for example,an increased coolant requirement exists because of a high operatingtemperature or because of special configurations of the internal coolingsystem in a gas turbine blade/vane, which configurations require highcoolant pressures or exhibit a high leakage rate.

In the case of different coolant requirements at different locations inthe internal cooling system, it is advantageous for a plurality ofsupply ducts to be present with different cross sections and transitioncross sections of the inlet openings, which transition cross sectionsare respectively matched to the differ-ent cross sections. In this way,the coolant can be individually matched to the coolant requirements ofthe different regions of the gas turbine blade/vane. By this, thecoolant consumption is reduced to the necessary extent. The manufactureof the supply ducts of different sizes or the manufacture of crosssections of different sizes is possible in one casting manufacturingprocess. It is only necessary to match the diameter of the core rib tothis requirement.

In order to obtain a large distribution space with reduced flow eddies,it is advantageous for the lowest longitudinal rib of the blade root,which is nearest to the axis of rotation of the gas turbine, to extendalong a principal axis of the gas turbine blade/vane. The blade root isheld, by way of its longitudinal ribs, on undercuts of the disk grooveinto which it is inserted. The distribution space for the cooling mediumis accommodated in the lowest longitudinal rib. In order to achieve thelargest possible distribution space, and therefore little eddying of thecoolant, the blade root according to the invention is lengthened in theregion of the lowest longitudinal rib. This lengthening takes placealong the principal axis of the gas turbine blade/vane i.e. at rightangles to the periphery of the disk when the gas turbine blade/vane hasbeen inserted. Due to the lengthened configuration of the lowerlongitudinal rib, the stability of the holding appliance in the bladeroot is further ensured and the rib can be easily lengthened in themanufacturing process of the gas turbine blade/vane by the core root ofthe casting core having a thicker configuration.

For the manufacture of the gas turbine blade/vane, in particular for thesubsequent machining of the blade root and for ensuring adequatestability, it is advantageous for the inlet openings of the supply ductsto be located at the level of the transition flank between the lowestlongitudinal rib and the longitudinal rib located above it. This ensuresthat the region of the distribution space is only surrounded by thelowest longitudinal rib. A transition flank, whose slope ensures secureholding of the blade root of the gas turbine blade/vane in the undercutof the disk is provided, is located between two longitudinal ribs. Thearrangement proposed for the inlet openings of the supply ducts ensuresthat subsequent work on the blade root after the casting process cantake place in a defined region without the blade/vane being damaged, theregion of the distribution space being located in each case within thelowest longitudinal rib. The lengthening of the longitudinal rib cantherefore be almost arbitrarily adjusted.

The object directed toward a casting appliance for the manufacture of agas turbine blade/vane with a distribution space is achieved, forexample, by an appliance for casting a gas turbine blade/vane having acasting core. It further preferably has core ribs forming the supplyducts, the casting core having a core root forming the distributionspace. The core ribs are preferably formed in one piece with that coreroot, with a continuous transition being present from the core root tothe core ribs.

In addition to an outer shell, the casting appliance has an innercasting core. The casting core is used when casting the gas turbineblade/vane in order to keep a predetermined, inner region of the gasturbine blade/vane free from cast material. The region kept freecomprises the inner cooling system, the supply ducts and thedistribution space. The supply ducts are kept free by elongatedextensions of the casting core, the so-called core ribs. Thedistribution space is formed by a region which is widened relative tothe core ribs and has a certain thickness and height, the so-called coreroot. The core root is configured in one piece with the core ribs. Theone-piece configuration of the two parts of the casting core permits arounded configuration of the transition between the supply ducts and thedistribution space.

The rounded configuration of the transition between the supply ducts andthe distribution space always occurs in the same manner, as specified bythe shape of the casting core. This permits exact maintenance ofpredetermined dimensions. It permits desired dimensions of the internalcooling system of the gas turbine blade/vane to be ensured in such a waythat they can be reproducibly adjusted for a complete series of gasturbine blades. This provides a basis for a low-cost and reliablemanufacture of internally cooled gas turbine blades/vanes.

Because the casting core is configured in one piece, it is very stablewith respect to the deformation forces which appear due to thesolidification of the melt.

The transition from the core root to the core ribs is designed, in sucha way that it takes place continuously in each case, by the crosssection being preferably continuously increased from the core ribs tothe core root. Because of the continuous transition from the core ribsinto the core root, no reworking of the inlet openings of the supplyducts is necessary after the casting process in order to ensure a lowflow resistance. This correspondingly dispenses with one operationalstep in the manufacture of the gas turbine blade/vane.

It is advantageous for the core ribs to merge into the core root withincreasing cross section, which core root has a thickness which islarger than the thickness of the core ribs. This permits an additionalsubstantial reduction in the flow resistance of the coolant flow.

A further improvement to the flow properties of the transition from thedistribution space to the supply ducts is provided by the rounded coreribs running out into a curved surface which ends in the core root. Thissurface forms a throat, which is provided before the actual inlets tothe supply ducts and which supports a continuous and low-eddy deflectionof the coolant flow into the supply ducts. In addition, such a castingcore is simpler to manufacture and can also be calculated moresatisfactorily with respect to its flow properties.

The task directed toward a method of manufacturing a gas turbineblade/vane, using an appliance as described for the casting, is achievedby the distribution space and the supply ducts being cast using theone-piece casting core.

Due to the use of the one-piece casting core, the casting process isdimensionally more accurate and, at the same time, less time-consuming,because the individual parts of the casting core can be installedjointly. With this method, the distribution space no longer needs to besubsequently recessed mechanically. This complicated measure, which hasessentially to be carried out manually, represents a time-consuming andcostly step in the manufacture of a gas turbine blade/vane withdistribution space. The use of the one-piece casting core, as proposed,now makes this process superfluous. In addition, the dimensions, andtherefore the coolant flow through the inlet openings of the supplyducts and the distribution space, can be reproducibly adjusted.

If, however, further changes to the distribution space and/or to theinlet openings of the supply ducts are still necessary and/or desirable,the distribution space can be mechanically reworked as a supplementarymeasure. As compared with the usual mechanical working, this issimplified by the fact that the major part of the material to be removedis already lacking due to the casting process. Only small corrections,which involve little manufacturing complication, have therefore to becarried out.

BRIEF DESCRIPTION OF THE DRAWINGS

The gas turbine blade/vane, the appliance and the method formanufacturing the gas turbine blade/vane with a cast distribution spaceare explained in more detail using the embodiment examples shown in thedrawings. In these:

FIG. 1 shows, in perspective side view, an excerpt from the disk and theblade/vane root,

FIG. 2 shows a perspective view from below onto the blade/vane root andthe distribution space,

FIG. 3 shows a view from below onto the distribution space, the inletopenings and the supply ducts,

FIG. 4a shows a longitudinal section through a disk feed duct, thedistribution space and the supply ducts of the blade/vane root of FIG.3,

FIG. 4b shows a cross section through the distribution space of FIG. 3,

FIG. 5 shows a perspective side view of the lower part of the castingcore and

FIG. 6 shows a cross section through the core rib and the core root ofFIG. 5.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

FIG. 1 shows, diagrammatically and not to scale, a construction inprinciple of the root region of a gas turbine blade/vane 1, inserted ina disk 3 of a gas turbine. The disk 3 can be rotated about therotational axis 14 of the gas turbine. The gas turbine blade/vane 1 isheld by way of its blade/vane root 2, which has two longitudinal ribs13, 13′, in a disk transverse groove 60 of the disk 3. The blade/vaneroot 2 is supported on undercuts 12 of the disk 3, by way of itslongitudinal ribs 13, 13′, against the centrifugal forces actingparallel to the longitudinal direction 15 of the gas turbine blade/vane15 when the disk 3 is rotating about the axis of rotation 14.

The disk 3 has a feed duct 6 and the blade/vane root 2 has a pluralityof supply ducts 4, which are in flow connection with one another bymeans of a distribution space 5. Coolant 80 can be fed from the disk 3into the internal cooling system of the gas turbine blade/vane 1 bymeans of this passage system. The coolant 80 is preferably cooling air.The distribution space 5 exhibits rounded or beveled inlet openings 7 ofthe supply ducts 4. In this way, the feed through coolant 80 is fed withminimum flow losses to the internal cooling system through thedistribution space 5 and into the supply ducts 4.

At its base 70, the distribution space 5 is open toward the feed duct 6.Practically no flow losses, therefore, occur at this base 70. Thedistribution space 5 is rounded in the manner of an ellipsoid. In itscross-sectional shape parallel to its base 70, it has a shape of anellipse which is contracting. In the cross-sectional area 9 at rightangles to this, shown in FIG. 4b, it has the cross-sectional shape of asemi-ellipse with continuously changing cross section. Thissemi-elliptical shape is interrupted by the rounded inlet openings 7 ofthe supply ducts 4. The transitions between the inlet openings 7 of thesupply ducts 4 and the semi-ellipse of the distribution space 5 have arounded design so that they form no flow resistance worth mentioning. Inthis arrangement, the inlet openings 7 can be located both directlyadjacent to one another, therefore meeting one another, or can beadjacent to one another.

The regions between the inlet openings 7 of the supply ducts 4 arerounded so as to optimize the flow, i.e. there are no edges present.This also applies to the cross sections 8 of the feed duct 6 in the disk3 of the gas turbine. The cross section 8 of the feed duct 6 ispreferably matched to the local changes in the cross sections 9 of thedistribution space 5 at right angles to its base plane 70, as are theinlet openings 7 with the downstream cross sections 10. In this way, acoolant flow 80 necessary for cooling the remote regions of the gasturbine blade/vane 1 can be reliably set. The supply ducts 4 bound thedistribution space 5 by way of different cross sections 10 andtransition cross sections 11 respectively matched to them, and merginginto the distribution space 5. In this way, a coolant flow 80 ofdifferent strengths, which depends in each case on the cross section 10of the supply duct 4, can be introduced into a predetermined region ofthe internal cooling system. This permits individual matching of thecooling.

The gas turbine blade/vane 1, which is represented in FIG. 1, ismanufactured in a single casting process, the distribution space 5 beingformed by a casting core 18 with the core ribs 19, which keep the supplyducts 4 free of cast material. The distribution space 5 has a height 90,which approximately agrees with the height 16 of the distance from thelower part of the lower longitudinal rib 13 to the transition into thefollowing longitudinal rib 13′ of the blade/vane root 2. In order toobtain a large distribution space 5 with the lowest possible flowresistance, it is correspondingly advantageous for the lowerlongitudinal rib 13 to be lengthened along a principal axis 15 of thegas turbine blade/vane 1. In the case of a distribution space 5 enlargedin this way, only a small proportion of eddying of the coolant flow 80is to be found within the distribution space 5 and on transition intothe inlet openings 7.

FIG. 2 shows a plan view onto the base 70 of the blade root 2, inperspective view. Rounded and/or beveled inlet openings 7 of the supplyducts 4 emerge from the distribution space 5. The longitudinal ribs 13,13′ are configured with undercuts 12.

FIG. 3 shows a direct view onto the lower surface of the blade root 2.The supply ducts 4 have an oval or elliptical shape, which isparticularly favorable to the flow. The inlet openings 7 are also,correspondingly, elliptically matched, the cross section of theelliptical inlet openings 7 continuously decreasing from thedistribution space 5 to the supply ducts 4.

FIG. 4a shows a longitudinal section through blade/vane root 2 and disk3. The coolant flow 80 passes from the feed duct 6, with diameter 8,into the distribution space 5 and, through the inlet openings 7, intothe supply ducts 4. The coolant flow 80 is fed unhindered into theinternal cooling system of the gas turbine blade/vane 1 through therounded inlet openings 7 and the rounded distribution space 5 andlikewise through the rounded opening 110 of the feed duct 6. Thedistribution space 5 has a maximum height 90.

FIG. 4b shows a cross section through the view of FIG. 3. The blade/vaneroot 2 of the gas turbine blade/vane, which is intersected by thedistribution space 5, is shown. The distribution space has an ellipticalcross section with the cross-sectional area 9.

FIG. 5 shows a casting core 18, which represents the essentialconstituent part of the appliance for casting a gas turbine blade/vane1. The casting core 18 has core ribs 19 and a core root 20. The coreribs 19, with the thickness 21, form the supply ducts 4 of the gasturbine blade/vane 1 during the casting operation. The core root 20 andthe core ribs 19 have a one-piece configuration and the core ribs 19merge with increasing cross section 21 into the core root 20. Thistransition takes place in a continuously increasing cross section 21 sothat no step-changes to the thickness occur. The core ribs 19 arerounded and preferably run out into a curved surface 24 which ends inthe core root 20. In this way, the distribution space 5 has a shapeafter the casting operation which is particularly favorable to the flow.FIG. 6 shows, in a longitudinal section through the core root 20 and acore rib 19, the continuous transition of the thickness 23 of the corerib 19 into the thickness 22 of the core root 20.

A casting core 18, as described above, is used during the manufacture ofthe gas turbine blade/vane 1 described above. It permits simplemanufacture both of a large distribution space 5 and of a continuoustransition from the distribution space 5 to the supply ducts 4 of thegas turbine blade/vane without reworking of the gas turbine blade/vane 1being necessary in this region. It is, however, readily possible tomechanically rework such a cast gas turbine blade/vane 1 in itsdistribution space 5 in order, for example, subsequently to adapt thegas turbine blade/vane 1 to changed requirements or to use the samecasting core 18 for different models. In this case, an essential part ofthe material to be removed is already kept free by the core root 20. Thesubsequent mechanical working is therefore only a correction, which canbe carried out rapidly and at low cost.

The invention being thus described, it will be obvious that the same maybe varied in many ways. Such variations are not to be regarded as adeparture from the spirit and scope of the invention, and all suchmodifications as would be obvious to one skilled in the art are intendedto be included within the scope of the following claims.

What is claimed is:
 1. A gas turbine blade through which coolant of aninternal cooling system flows, comprising: a blade root inserted into arotatable disk of the gas turbine; a plurality of supply ducts; and adistribution space, wherein coolant can be fed to the supply ductsthrough a feed duct of the disk, the feed duct communicating with thesupply ducts via the distribution space, and wherein the distributionspace is a cast distribution space including at least one of rounded andbeveled inlet openings for the supply ducts.
 2. The gas turbine blade asclaimed in claim 1, wherein the distribution space is rounded in themanner of an ellipsoid.
 3. The gas turbine blade as claimed in claim 2,wherein the inlet openings meet one another in a manner which optimizesthe flow.
 4. The gas turbine blade as claimed in claim 2, wherein theinlet openings are adjacent to one another in a manner which optimizesthe flow.
 5. The gas turbine blade as claimed in claim 1, wherein theinlet openings meet one another in a manner which optimizes the flow. 6.The gas turbine blade as claimed in claim 1, wherein a cross section ofthe feed duct and local changes to cross sections of the distributionspace are matched to the cross sections of the outlet openings locateddownstream.
 7. The gas turbine blade as claimed in claim 6, wherein aplurality of supply ducts are present with different cross sections andtransition cross sections of the inlet openings, respectively matched tothe different cross sections.
 8. The gas turbine blade as claimed inclaim 1, wherein a plurality of supply ducts are present with differentcross sections and transition cross sections of the inlet openings,respectively matched to the different cross sections.
 9. The gas turbineblade as claimed in claim 1, wherein the blade root includeslongitudinal ribs, which engage in undercuts on the disk and of which arelatively lowest, which is relatively nearest to an axis of rotation ofthe gas turbine, is lengthened along a principal axis of the gas turbineblade.
 10. The gas turbine blade as claimed in claim 9, wherein theinlet openings of the supply ducts are located at a level of atransition flank between a relatively lowest longitudinal rib and a nextrelatively lowest longitudinal rib located above it.
 11. The gas turbineblade as claimed in claim 1, wherein the inlet openings of the supplyducts are located at a level of a transition flank between a relativelylowest longitudinal rib and a next relatively lowest longitudinal riblocated above it.
 12. An appliance for casting the gas turbine blade asclaimed in claim 1 comprising: a casting core, including core ribsforming the supply ducts, and a core root forming the distributionspace, the core ribs being formed in one piece with the core root, and acontinuous transition being present from the core root to the core ribs.13. The appliance for casting a gas turbine blade as claimed in claim12, wherein the core ribs merge into the core root with increasing crosssection, the core root having a thickness which is relatively largerthan the thickness of the core ribs.
 14. The appliance for casting a gasturbine blade as claimed in claim 13, wherein rounded core ribs run outinto a curved surface which ends in the core root.
 15. A method formanufacturing a gas turbine blade using the appliance as claimed inclaim 13, comprising: casting the distribution space and the supplyducts using the casting core.
 16. The appliance for casting a gasturbine blade as claimed in claim 12, wherein rounded core ribs run outinto a curved surface which ends in the core root.
 17. A method formanufacturing a gas turbine blade using the appliance as claimed inclaim 16, comprising: casting the distribution space and the supplyducts using the casting core.
 18. A method for manufacturing a gasturbine blade using the appliance as claimed in claim 12, comprising:casting the distribution space and the supply ducts using the castingcore.
 19. The gas turbine blade as claimed in claim 1, wherein the inletopenings are adjacent to one another in a manner which optimizes theflow.